This invention relates generally to gas turbine engines, and, more specifically, to turbine airfoil cooling.
Efficiency is a primary concern in the design of any gas turbine engine. Historically, one of the principle techniques for increasing efficiency has been to increase the gas path temperatures within the engine. Using internally cooled components made from high temperature capacity alloys has accommodated the increased temperatures. Turbine stator vanes and blades, for example, are typically cooled using compressor air. Cooling is typically extracted from the compressor at a temperature lower and pressure higher than the core gas passing through the turbine section. The cooler compressor air provides the medium for heat transfer and the difference in pressure provides the energy required to pass the cooling air through the stator or rotor stage. It will be understood that compressor bleed air for such cooling will be unavailable to support combustion in the combustor. A significant percentage of the work imparted to the air bled from the compressor, however, is lost during the cooling process. The lost work does not add to the thrust of the engine and negatively effects the overall efficiency of the engine. A person of skill in the art will recognize therefore, that there is a tension between the efficiency gained from higher core gas path temperatures and the concomitant need to cool turbine components and the efficiency lost from bleeding air to perform that cooling. There is, accordingly, great value in maximizing the cooling efficiency of whatever cooling air is used.
Thus, to minimize any sacrifice in engine performance due to the unavailability of cooling airflow to support combustion, any scheme for cooling blades and vanes must optimize the utilization of compressor bleed cooling air. Airfoil cooling is accomplished by external film cooling, internal air impingement and forced convection, either separately or a combination of all cooling methods.
In forced convection cooling, compressor bleed air flows through the internal cavities of the blades and vanes, continuously removing heat therefrom. Typically, compressor bleed air enters internal cavities of the blades and vanes through one or more inlets which discharges into the internal cavities.
Film cooling has been shown to be very effective but requires a great deal of fluid flow to be bled off the compressor for cooling. Further, film cooling is actively controlled in a complex and expensive manner. Also, the fabrication and machining of an airfoil with film cooling holes not only adds a degree of complexity but is also costly. It will also be appreciated that once the cooling air exits the internal cavity of the airfoil and mixes with the hot combustion gases, a severe performance penalty is incurred due to the mixing process and the different temperature levels of the mixing flows.
In many cases, it is desirable to establish a film of cooling air along the surface of the stator or rotor airfoil by bleeding cooling air out of cooling holes. The term “bleeding” reflects the small difference in pressure motivating the cooling air out of the internal cavity of the airfoil. The film of cooling air traveling along the surface of the airfoil directs the flow of high thermal energy hot gas away from the airfoil, increases the uniformity of the cooling, and thermally insulates the airfoil from the passing hot gas stream flow. A person of skill in the art will recognize, however, that film cooling is difficult to establish and maintain in the turbulent environment of a gas turbine.
A known method of establishing film cooling involves positioning cooling holes in or adjacent the leading edge of an airfoil in a “showerhead” arrangement. The showerhead typically includes a row of cooling holes on either side of the leading edge. The cooling holes are angled aft and are often diffused to facilitate film formation. In some cases, the showerhead includes a row of holes positioned directly on the leading edge. U.S. Pat. No. 5,374,162 discloses an example of such an arrangement.
One problem associated with using holes to create a cooling air film is the film's sensitivity to pressure difference across the holes. Too great a pressure difference across a cooling hole will cause the air to jet out into the passing core gas rather than aid in film formation. Too small a pressure difference will result in negligible cooling air flow through the hole, or worse, an in-flow of hot core gas. Both cases adversely affect film cooling effectiveness. Another problem associated with using holes to establish film cooling is that cooling air is dispensed from discrete points along the span of the airfoil, rather than uniformly and along a continuous line. The gaps between cooling holes, and areas immediately downstream of those gaps, are exposed to less cooling air than are the holes and the spaces immediately downstream of the holes, and are therefore more susceptible to thermal distress. Yet another problem associated with using holes to establish film cooling is the stress concentrations that accompany each hole. Stress concentrations develop when loads (typically resulting from dynamic forces or thermal expansion) are carried by narrow expanses of material extending between adjacent holes. Film cooling effectiveness generally increases when the cooling holes are closely packed and skewed aft at a shallow angle relative to the external surface of the airfoil. Skewed, closely packed apertures, however, are more prone to stress concentrations. Thus, film cooling requires a greater amount of cooling air with the possibility of inadequate cooling of the outer surfaces of the airfoil.
Some prior art configurations have cooling holes disposed in the leading edge aligned with an average stagnation line, that extend perpendicular to the external surface of the airfoil. High temperature core gas (which include air and combustion products) encountering the leading edge of an airfoil will diverge around the suction and pressure side portions of the airfoil, with some of the gas impinging on the leading edge. The point along the airfoil where the velocity of the core gas flow decelerates to zero (i.e., the impingement point) is referred to as the stagnation point. There is a stagnation point at every spanwise position along the leading edge, and collectively those points are referred to as the stagnation line. Air impinging on or adjacent the leading edge is subsequently diverted around either side of the airfoil. In actual practice, rotor speeds and core gas velocities vary depending upon engine operating conditions as a function of time and position along the leading edge. Such a cooling hole arrangements can experience an asymmetrical cooling air distribution. For example, an actual stagnation line shift to one side of a row of cooling holes can urge exiting cooling air to one side of the row, consequently leaving the opposite side starved of cooling air. The fact that the stagnation line can and does shift during airfoil operation illustrates that locating holes on the average stagnation line will not remedy all cooling air distribution problems. Cooling holes extending perpendicular to the external surface and skewed spanwise do not resolve the potential for asymmetrical cooling air distribution.
Also, some prior an configurations employ a trench at the leading edge with cooling holes exiting into a trench. The cooling holes are discrete cooling points with uncooled areas inbetween. The cooling holes must fill the trench such that the cooling air can dwell within the trench and bled out of the trench. Key to use of the trench requires that the trench be filled with cooling air. However, the addition of too many cooling holes requires a significant amount of cooling air from the compressor that can negatively impact turbine efficiency. Also, too many cooling holes located at the leading edge of the airfoil can create undesirable thermally induced stresses in the metal between the holes. Further, if the trench is not adequately filled then there is a risk of uncooled areas between the holes. Finally, the prior art configurations rely on film cooling to cool the leading edge and aft of the leading edge thus requiring a significant amount of cooling air to ensure adequate film coverage.
Turbine engine blade designers and engineers are constantly striving to develop more efficient ways of cooling airfoils and prolong turbine blade life and reduce engine operating cost. Cooling air used to accomplish this is expensive in terms of overall fuel consumption. Thus, more effective and efficient use of available cooling air in carrying out cooling of turbine airfoil and, in particular, a leading edge of an airfoil is desirable, not only to prolong turbine airfoil life, but also to improve the efficiency of the engine as well, thereby lowering engine operating cost. Consequently, there is a continuing need for airfoil cooling designs that will make more effective and efficient use of available cooling air.
Thus, what is needed to extend the durability of a turbine airfoils is an improved cooling design suitable for use at a leading edge of an airfoil that provides reliable, complete and uniform film cooling while optimizing the cooling air necessary and reducing the stress associated with the spacing between the cooling apertures. Further, what is needed is a leading edge cooling configuration that employs film cooling, impingement cooling and convective cooling and that can be manufactured with the casting process.